Gas turbine propulsion systems for aircraft must deliver high performance in a compact, lightweight configuration. This is particularly important in lower power propulsion systems. In these applications, it is advantageous to utilize “axi-centrifugal” compressors, which includes one or more axial compressor stages followed by a centrifugal stage. While all-axial compressors may be developed for smaller engines, the last blade rows of the compressor have airfoils that are very small and, thus, highly sensitive to destabilizing features like the clearance gap between the blade tip and the outer case. The centrifugal stage of an axi-centrifugal compressor is less sensitive to these factors and, therefore, provides higher performance that is easier to retain. With an understanding of the benefit of using an axi-centrifugal compressor in these applications, it is important to evaluate modification of other components in the gas turbine propulsion system, such as the fan in combination with the axi-centrifugal compressor (i.e., the compression system).
In general, fans for aviation propulsion gas turbine engines must meet stringent durability criteria while delivering high performance and low weight at an acceptable cost to the commercial industry. Larger thrust engines inherently benefit from the larger size of the fan blades to meet bird and ice ingestion criteria of the Federal Aviation Administration. Larger thrust engines (i.e. greater than 15 klbf Sea-Level Takeoff (SLTO) thrust) also tend to exhibit increased bypass ratios and require less fan pressure rise which allows them to run at fan tip speeds as low as 1100 ft./sec and, thus further enhance their ability to meet these ingestion criteria.
The fan must be free of any aerodynamic or aero-elastic instability such as surge and flutter throughout the required operating regime. These requirements have historically driven aviation propulsion engines to one of two solutions. The first solution incorporates individual fan blades or airfoils mechanically inserted into a separate disk. Fans of this type typically have a fan blade thickness distribution resulting in a “1/rev” design, wherein the fundamental vibratory mode of the fan blade is above the first engine order at the fan maximum rotational speed. The interface between these airfoils and the disk introduces mechanical damping when the fan blades undergo movement relative to the disk, which tends to mitigate potential failure modes resulting from an aero-elastic instability known as flutter.
The second solution provides an integrally bladed disk fan also known as a blisk fan, which have blades integrally formed with or materially joined to a disk. For blisk fans, the distribution of fan blade thickness is modified to obtain a “2/rev” design or greater such that the fundamental vibratory mode of the fan blade is above the second engine order at the fan maximum rotational speed to address aeroelastic and aero-mechanical vibration and instability. By nature, a blisk fan is more prone to flutter because the fan blade is material joined to the rotor hub, thus reducing the inherent dampening in the fan. Increasing the fundamental modal frequency by increasing blade thickness in the airfoil region nearest the disk can significantly mitigate flutter even in a blisk fan. This thickness increase is often incorporated along with an increase in the chord of the blade, resulting in a decrease in aspect ratio. While this can be effective for reducing flutter, it often results in an unacceptable increase in weight. While the fan blades may be hollowed to reduce weight, the added cost associated with manufacture introduces an unacceptable increase in cost to the commercial industry.
Hence, there is a need for a simple and effective blisk fan configuration for use with an axi-centrifugal compressor in lower powered gas turbine propulsion systems, which meets certain performance requirements in a compact durable and cost-effective design.